An aircraft gas turbine engine or jet engine draws in and compresses air with an axial flow compressor, mixes the compressed air with fuel, bums the mixture, and expels the combustion product through an axial flow turbine section that powers the compressor. The turbine section of the engine includes one or more disks, each disk including a plurality of blades projecting from its periphery. The hot exhaust gases strike the blades causing the disk(s) to rotate. The rotating disk(s) are attached to a shaft that also drives a compressor. The compressor is also made from rotating disks, each disk having a plurality of blades projecting from its periphery. The disk turns rapidly on a shaft as the shaft is rotated by the turbine, and the curved blades draw in and compress air in somewhat the same manner as an electric fan.
The turbine blades on the disk are in the hot exhaust gases resulting from the combustion of the fuel and rotate at very high speeds. Thus the blades operate in an oxidative and corrosive environment, and are subjected to high operating stresses. In order to survive these harsh conditions, the turbine blades are made from superalloys, an expensive blend of elements that provide oxidation resistance, corrosion resistance and strength. These superalloys are further strengthened in preferred directions by various mechanisms which include growing the turbine blades as directional grains or even as single crystals.
The superalloys used for turbine blades include nickel-based superalloys, iron-based superalloys and cobalt-based superalloys. These superalloys can be further strengthened by precipitation mechanisms. For example, gamma prime (γ′) phases comprising Ni3Al are precipitated in the gamma matrix of the FCC crystal structure of the alloy by appropriate solutioning and aging treatments. Controlling the gamma prime phases, both the size and distribution for these as-cast, new parts is well-known. The turbine blade comprises an airfoil portion that extends into a hot gas stream, a dovetail portion that attaches the blade to the turbine disk and optionally a platform portion that separates the airfoil portion from the dovetail portion. A shank portion is intermediate the platform portion and the dovetail portion. The turbine blades are also provided with environmental coatings and/or thermal barrier coatings to further improve their survivability in the hot, corrosive, oxidative environment of a turbine engine.
The turbine blades nevertheless are subject to damage as a result of operation in the gas turbine engine. This damage can be both mechanical in nature as well as metallurgical in nature. The turbine blades are expensive to produce, so that it is desirable from an economic standpoint to repair the blades rather than replace them whenever possible. In many situations, the blades can be repaired by removing any remaining protective coatings, followed by welding damaged mechanical areas and reworking the weld repaired areas to restore the dimensions as required, followed by reapplication of the protective coatings.
The metallurgical damage is inherent as a result of normal operation of the gas turbine engine. The blade material, typically a high temperature superalloy, derives its corrosion and oxidation resistance by selection of a careful combination of elements. Its strength is a result of formation of fine precipitates by precipitation hardening. However, operation of the blades at the high temperatures of the turbine engine for extended periods of time results in fatigue, creep-rupture and rafting. Rafting, when it is present, occurs in the turbine blade from the platform to the top, and more specifically, in the blade trailing edge below the tip and above the platform. As used herein, rafting means coarsening of the γ′ precipitate and precipitate alignment in the direction of applied stress. This region having the coarsened γ′ is characterized by reduced strength.
Welding of the blades to repair mechanical damage is known. The damage repaired by welding is not limited to blades removed from engine service, but can also be new blades requiring repair of casting defects or damage resulting from testing or machining. Although various weld repair methods are available, one illustrative weld repair method is SWET welding, that is, Superalloy Welding at Elevated Temperatures. This process was developed by General Electric Aircraft Engines in the 1980's to commercially repair damaged turbine blades refurbished for its airline customers. The process entails heating of the blade airfoil to an elevated temperature using a heat source such as quartz lamps or induction coils. The heat source is focused in a narrow region of the blade undergoing repair, typically at the blade tip, and heat is localized at the blade tip. An area of the blade to be repaired is then welded while maintaining the blade at elevated temperature. A variation of SWET welding techniques appears in various prior art references.
The problem with welding of blades is that the weld area and heat affected zone (the “HAZ”) are heated above the solutioning temperature of the gamma prime. As the blade cools, heat is conducted away from the narrow area of weld repair and the HAZ, but γ′ will precipitate in this area. As heat is transferred away from the repair area, and in portions of the HAZ, the temperature of the metal is increased, but not to a temperature sufficient to raise the alloy above the solutioning temperature of the γ′. While fine γ′ precipitate forms in the narrow region of weld repair and a portion of the HAZ, the γ′ precipitate in narrow adjoining regions further coarsens and is characterized by further reduced strength. Thus, mechanical repair by welding does not provide a solution to the metallurgical problems related to extended use of precipitation-hardened alloys at elevated temperatures, and in certain cases, may further exacerbate the problem.
Although the prior art discloses that the weld area is desirably stress relieved and rapidly cooled to a temperature below the γ′ precipitation hardening temperature, the prior art does not address the problem of rafting in other regions of the airfoil portion of the blade that may be distal weld repair. The prior art references indicate that, in order to reduce the residual stresses in the repaired article, the weld-repaired blade (i.e. the entire blade) is placed in a fixture and stress relieved in a furnace, as is typically done on weld-repaired articles subject to stress, although the written description is otherwise devoid of a discussion of how post-weld treatments are accomplished. The prior art does not recognize the need to restore the metallurgical properties of airfoils not subject to weld repair.
A method for sintering a wear resistant layer to a blade tip is disclosed in U.S. Pat. No. 4,818,833 issued Apr. 4, 1989. This patent identifies utilizing a radiant heat source to heat the blade tip. The radiant heat source is a graphite susceptor having heating chambers which extend into the susceptor. No post-sintering heat treatments are discussed; however the patent does disclose coarsening of the gamma prime near the blade tip that was judged to be acceptable, but does not recognize the problem of continued coarsening of this gamma prime that will inevitably result from high temperature exposures for long periods of time. Of course, obvious methods of heating may be substituted for the radiant heat source, such as for example, induction coils, but such substitutions do not provide recognition of the problem of continued coarsening, or alleviation of coarsening that has already occurred.
U.S. Pat. No. 6,020,571 (the '571 patent) and U.S. Pat. No. 6,124,568 (the '568 patent), assigned to the assignee of the present invention and incorporated herein by reference in their entirety disclose acceptable methods of welding of precipitation hardenable nickel-based superalloy blades by carefully controlling heat input during the welding process. While effective for advancing the art of welding, the '568 patent only deals with the problem of γ′ precipitation resulting from the welding operation, and this problem is controlled by controlling the heat input in the localized region undergoing repair in accordance with a desired temperature profile.
In order to achieve a uniform γ′ precipitate structure throughout the blade, either as a result of post-welding coarsening or as a result of rafting due to extended exposure to high temperatures, one current practice is to solution the entire blade, thereby dissolving all γ′, followed by reprecipitation of the γ′. This process is effective in providing uniform γ′ in the airfoil portion of the blade, but presents other problems. Specifically, dovetail regions of the blade and the shank regions of the blade, both located below the platform region, are subjected to high stresses, either from machining operations or from service induced stresses as a result of being positioned in the disk at high rotational speeds, or both. These operations result in residual stresses. The portions of the blade below the platform also experience an operating temperature significantly lower than the airfoil portion of the blade, which extends into the hot gas stream. Thus, a high temperature solutioning treatment of those portions of the blade below the platform is undesirable as it can result in recrystallization of these portions due to the residual stresses. Since many modern blades are either directionally solidified (providing large columnar grains oriented parallel to the longitudinal axis of the blade) or are solidified as single crystals, recrystallization in this region is undesirable as it reduces strength of the blade in this region, which can be a limiting factor for the mechanical properties of the entire blade.
What is needed is a technique that permits restoration of metallurgical properties of the airfoil portion of a blade having γ′ precipitates to eliminate the problem of rafting. The rafting can be a result of extended use of the blade at elevated temperatures and load. The restoration of the metallurgical properties of the airfoil portion of the blade should be accomplished without adversely affecting the as-cast grain structure of the portions of the blade at or below the blade platform, such as the dovetail area of the blade.